The present invention relates to the field of gas turbines, in particular those to be found in turbomachines, and by way of non-limiting examples but not only in the turbine engines of helicopters or in the turbojets for airplanes.
The present invention relates more particularly to the compression stage of such gas turbines that constitute the main power plant of an aircraft.
Still more precisely, the present invention relates to a centrifugal compressor of a turbine engine, the compressor comprising:                a cover including an upstream end and a downstream end;        a casing presenting an upstream edge and a downstream edge; and        a bladed impeller mounted to rotate in said casing;        
said cover being designed to cover the blades of the impeller so as to define an outside surface of a gas-flow passage extending between the upstream and downstream edges of the casing, being fastened to the upstream edge of the casing via its upstream end while its downstream end remains free.
Conventionally, the compressor is placed between a fresh air inlet and a combustion chamber, the role of the compressor being to compress the fresh air entering into the gas turbine and to convey the compressed air into the combustion chamber in order to be mixed with fuel.
Furthermore, it is known that an impeller comprises a plurality of blades extending generally radially from an impeller hub, which hub is fastened to a rotary shaft of the gas turbine.
Thus, the gas stream initially enters into the casing of the compressor via an upstream inlet, and then flows along a gas-flow passage defined between an outside surface defined by the cover and an inside surface defined by a surface of the impeller hub, while being compressed and driven in rotation about the axis of the impeller prior to being exhausted through a downstream outlet of the compressor, it being specified that the terms “upstream” and “downstream” are taken relative to the flow direction of the gas in the gas-flow passage through the compressor.
Generally, the stream of compressed gas leaving the impeller then penetrates into a diffuser prior to entering into the combustion chamber.
It can thus be understood that the cover defines the outside surface of the gas-flow passage, with the inside surface of the passage being formed by a surface of the impeller hub from which the blades extend.
In order to control the thermomechanical behavior of the cover, its downstream end is generally left free, i.e. it is not fastened to the downstream edge of the casing.
This configuration serves to avoid the cover being secured in a statitically overdetermined manner which would have the potential of damaging control over the clearances between the impeller and the cover.
Nevertheless, that solution is not perfect: certain degraded behaviors of the compressor, such as pumping or other unstable phenomena, for example, can appear and can lead to sudden variations of pressure within the impeller of the compressor.
Insofar as the downstream end of the cover is free, it will be understood that it can deform slightly as a result of pressure variations inside the compressor, and that such deformation might lead to the cover coming into contact with the blades of the impeller. When the pressure inside the compressor drops below that existing outside the cover, then the cover tends to deform so as to come into contact with the blades of the impeller. This deformation may also be due to vibration.
Naturally, it is extremely harmful both for the cover and for the impeller if the cover comes into contact with the blades of the impeller, where such contact might seriously damage the compressor.
Such a phenomenon may also occur when the gas turbine is being operated under extreme conditions.
One solution to the problem is to increase the clearance that exists between the cover and the blades of the impeller. Nevertheless, such a solution presents the drawback of reducing the efficiency of the compressor, and consequently of diminishing the performance of the gas turbine.